Blade elements for turbo machines



Oct. 3, 1961 A. R. HOWELL ET AL 3,002,675

BLADE ELEMENTS FOR TURBO MACHINES Filed Nov. 3, 1958 2 Sheets-Sheet 1 FIG. 4 i

Inven ars- Oct. 3, 1961 A. R. HOWELL ETA!- 3,

BLADE ELEMENTS FOR TURBO MACHINES 2 Sheets-Sheet 2 Filed Nov. 3, 1958 A ea tamzys 3,002,675 BLADE ELEMENTS FOR TURBO MACHINES Alun Raymond Howell, Cove, Farnborough, and Arthur Norman Smith, Waiton-on-Thames, England, assignors to Power Jets (Research and Development) Limited,

London, England, a British company Filed Nov. 3, 1958, Ser. No. 771,577 Claims priority, application Great Britain Nov. 7, 1957 Claims. (Cl. 230-116) This invention relates to the arrangement and mounting of blades in elastic fluid flow machines such as axial flow compressors, gas turbines or combinations thereof, of the kind having or providing at least two annular flow paths for fluid which extend concentrically with and coaxially of a rotor which has radial rotor blades extending into said flow paths thereby allowing energy transmission to occur between fluid in said flow paths and the rotor. Blades which thus extend into two annular flow paths are conveniently referred to as two tier blades and may be associated with the stator of the machine as well as with the rotor.

Hitherto, two tier blading has been proposed for rotors of machines of the kind described, in which the second tier blade, which extends within the outer of the two annular flow paths, is made integral with, and radially coextensive with, the first tier blade which, it will be understood, extends within the inner of the two annular flow paths.

In contradistinction to the aforesaid proposal, the present invention provides a two tier blading arrangement for elastic fluid flow machines in which the first tier blades carry peripheral shroud ring structure and second tier blades extend radially from spaced positions around said shroud ring structure intermediate the tips of the first tier blades.

Preferably, each first tier blade carries adjacent its tip portion a part of the peripheral shroud ring structure in the form of a segmental platform extending circumferentially on each side of the said tip portion by an amount equal to half the intended pitch of the first tier blades. Means may be provided at the terminal part of the circumferentially extended platform for supporting a root portion of second tier blade. This means may comprise a recess, or the like, which in co-operation with a similar recess in the platform of an adjacent blade forms a second tier blade root-retaining groove.

With a plurality of such first tier blades assembled side by side in machine structure, the segmental platform structure defines a peripherally continuous shroud ring constituting at least an annular part of the dividing wall between the two annular flow paths.

If desired the shroud ring may be an integral member and secured to the first tier blade tips.

In a further aspect, the invention comprises a gas turbine engine having an axially extending rotor and two concentric co-axial inner and outer gas flow paths, and a two tier blading arrangement associated with the rotor comprising a row of peripherally spaced first tier blades attached to the rotor and extending into the inner of the two flow paths, a shroud ring structure carried by tip portions of thefirst tier blades and defining at least a part of the dividing wall between said two gas flow paths and a series of second tier blades extending from posi tions along the shroud ring at positions thereon intermediate said first tier blades into the outer gas path.

In a still further and more particular aspect of the invention, there is provided a ducted fan type gas turbine engine incorporating an axially extending rotor, two coaxial and concentric inner and outer, annular flow paths for elastic fluid, a series of first tier turbine blades exam o tending into the inner of the two flow paths from peripherally spaced positions on the rotor for subjection to loading under elastic fluid flowing in that flow path, shroud ring structure carried by tip portions of said turbine blades, second tier fan blades extending from said shroud ring structure into the outer flow path for subjection to loading under elastic fluid flowing in that path, the second tier fan blades being mounted at such positions along the shroud ring as will, under loaded conditions, impose a couple on the turbine blades in opposition to that bending moment induced by the fan blades on the turbine blades.

One embodiment of the invention will now be described in its application to two tier blading for a rotor of a ducted fan engine with reference to the accompanying drawings in which FIGURE 1 is a diagrammatic view showing one half of a ducted fan gas turbine jet propulsion engine in axifl cross section which incorporates two tier blading according to the invention;

FIGURE 2 is a perspective view on an enlarged scale showing part of a row of two tier blading and its associated rotor;

FIGURE 3 is a plan view of a development of part of the blading shown in FIGURE 2, and

FIGURE 4 is a bending moment diagram showing an approximate comparison of the stress conditions due to bending which obtains between a conventional co-extensive two tier blade and staggered two tier blading shown in FIGURE 2 as applied to the engine shown in FIG- URE 1.

In FIGURE 1, the engine shown is generally of conventional layout and includes a main air intake 1 directing air in the direction of the arrow A along an axially extending flow path between a wall 2, supporting stator structure and an axial flow compressor rotor 3 into a combustion chamber 4. The latter is supplied with fuel by a fuel supply means 5 and combustion gases are delivered from the chamber 4 to a turbine including a rotor 6. At this region, the engine provides two annular, axially extending flow paths for fluid including an inner annular flowpath defined between a wall 7 carrying engine stator structure and the turbine rotor 6 and an outer concentric annular axially extending flowpath 8 defined between the wall 7 and a concentric annular wall 9. The flow path 8 has a forwardly facing inlet 10 for air and an axially spaced outlet 11. The turbine rotor 6 is mounted on a shaft 12 common to the compressor rotor 3 which is driven thereby. Both the inner and outer flow paths terminate in a common jet propulsion nozzle 11a.

A rotor 13 of a free turbine is supported downstream of the turbine rotor 6, the axis of the rotor 13 being coaxial with axis XX of the engine. The rotor 13 carries at its periphery a row of two tier blading. The latter provides first tier axial flow turbine blades 14 which extend, from the rotor 13, within the inner annular flow path, first tier blade tip shroud structure 15 of annular form extending circumfercntially within an annular gap 2a in the wall 7 and second tier axial flow fan blades 16 extending from the shroud structure 15 within the outer concentric flow path 8.

In operation, the hot gas eflluent from the combustion chamber 4 is firstly expanded through the blades of the turbine rotor 6 which drives the compressor rotor 3, and secondly expanded through first tier turbine blades 14 rotating rotor 13 of the free turbine in the opposite direction to the rotation of rotor 6. First tier turbine blades 14 thus drive the second tier fan blades 16 and a stream of air is drawn through inlet 10 compressed by fan blades 16 in the flow path 8 whence it is directed through outlet 11 as an annular air stream which combines with the hot gas efiluent from the first tier blades 14 to pass out through propulsion nozzle 11a. The free turbine rotor 13 and its associated blading thus acts as a thrust augmentor to augment the propulsive thrust of the engine.

The two tier blading as proposed hitherto has taken the form of elongated co-extensive blade elements. In contradistinction to these proposals, the two tier blading in this embodiment has its second tier blades staggered in relation to the first tier.

Reference to FIGURE 2 shows a part of a row of the two tier blading and the rotor 13. The first tier blades 14 are mounted in the rotor 13 by the usual root fixings at their attachment ends whilst blade stem portions 14a extend radially across the hot gas path and are accordingly formed by a heat resistant material, such as Nimonic (registered trade mark) alloy. The stem portion 14a of each blade 14 merges atits tip 14b with a segmental platform 15 of suitable curvature so as to form a segment of a continuous tip shroud ring structure within the annular gap 212 in the wall 7 where it carries stator structure and divides the concentric flow paths.

Each segmental platform 15 present a pair of opposite parallel side faces 15a which are inclined to the axis XX of the rotor such that planes containing the axis of the rotor and passing through the mid pitch position of the first tier blades 14 bisects the faces 15a. The bisection of a face 15a by such a plane is represented by the line PP in FIGURE 2.

As showneach face 15a is cut away as shown to form a recess 15b shaped as one half of a dovetail groove so that two such recesses on adjacent platforms 15 can cooperate to form a sing-1e dovetail shaped groove which is inclined lengthways with respect to the axis of the rotor 13. Each recess 15b includes a base portion 150 relieved as shown.

When two adjacent first tier blades 14. are assembled in the rotor 13 the adjacent faces 15a of a platform 15 on each blade abut one another to define an inclined dovetail groove whilst the platforms themselves co-opcrate to form a segmental part of tip blade shroud. Into the dovetail groove so, formed an attachment end of a second tier blade is inserted, the attachment end 21 being formed as a, dovetail root as shown. From the end 21 a stern portion 16a extends across the outer of the two coaxial fio-w paths. As shown in FIGURE 3 the stem portion is designed-to have a low stagger angle that is to say the angle 'y between the line ab parallel to the axis XX of the rotor and the line no tangential to and joining the leading edge and trailing edge of the blade is small, in this case about 25". In FIGURE 3 the directions of air flow and'the rotation of the rotor 13 are indicated by the arrows F and R respectively.

When a row of two tier blading 14, 16 is mounted in the rotor 13 in the manner described the platforms 15 co-operate to provide a continuous annular shroud ring structure at the tips of the first ti'er blades 14 and thereby to define a part of the dividing wall between inner and outer flow paths. The second tier blade serve to compress air in the outer flow path 8 when the blading is driven rotationally by hot gas flow through the first tier blading 14. as described above.

The advantageous stress condition obtained by the staggering of its' second tier fan blades as compared with the known co-extensive two tier blades having first tier turbine blades and second tier fan blades is demonstrated with reference to FIGURE 4 which shows bending moment diagrams of both co-extensi've blades and those arranged according to the embodiment described above with second tier blades at low stagger angle.

In FIGURE 4 the horizontal line OX represents the datum above and belowwhich bending moments of oppositesign are indicated. The vertical'lines OY andO Y represent the positions of the first tier blade root and junction of first and second tier blades, respectively, along the length of a two tier blade element.

Considering, firstly, the bending moment in a co-extensive two tier blade element, the work done by the second tier fan blade imposes a bending moment on the blade depicted by the line ABC (point'C representing the fan blade tip with zero bending moment). The bending moment due tothe gas load on the first tier turbine blade is the line D0 which is of opposite sign. The resultant bending moment on the first turbine tier blade would then correspond with the area OFBO the bending moment. at the point of attachment of the first tier blade root being indicated at OF.

Considering now the corresponding stress conditions due to bending on a part of a rotor blade row arranged in the manner of FIGURE-S. 2 and 3 and under load cone ditions characteristic in the operation of the engine shown in FIGURE 1, the line BC represents the bending moment on the second tier fan blade 16 due to work doneby that blade on the gas, the line D0 represents the moment on the first turbine tier blade 14. due to gas load and is of opposite sign as before. of the second tier blades, a bending moment G0 is imposed on the first tier blade 14 consequent upon the tangential force applied to the first tier blade from the fan blade. Additionally, however, the staggered relationship between the tiers means that the couple on the second tier fan blade due to the work it isdoing on the air is resisted by an equal and opposite couple on two adjacent first tier turbine blades. Thus the couple due to two adjacent second tier fan blades is effective to introduce an opposite couple in a first tier turbine blade located intermediate these two second tier blades. This opposite couple gives a bending moment HI to the first tier blade of opposite sign to the fan induced moment G01. Adding to the moment H! the moment CD0 on the first tier blade due to its own gas load gives a total'negative bending moment of 00 1K. Adding the latter total moment to the positive moment G0 algebraically, the resultant moment on the first tier blade will approximate to the dotted .line LI. As will be evident from FIGURE 4, this is considerably less than the moment PB on the first tier blade of a coextensive two tier blade. The magnitude of the respective bending moments on the first tier root attachment are proportional to the length OF for the co extensive blade and to the length LO for the staggered blade arrangement.

In the above example the second tier blade root retaining grooves are inclined to the axis of the rotor, these could be made helical and this may be preferable in some instances.

Straight axially extending dovetail grooves have the disadvantage that due to the diagonal disposition of the second tier blade on its dovetail shaped base or root portion, the distance between the stern portion of the blade where it adjoins the base and the inclined undercut portions of the dovetail groove will be greater in the case of axial extending dovetail, and hence the cantilevering effect on the undercut portions of'the dovetail will be greater, than in the case of helically formed axially inclined dovetail base and its complementarily shaped groove.

The employment of separate blades for the second tier permitsthese blades being made of a different material from that of the first tier blades. Thus in the recited example, where the rotor carries first tier turbine. rotor blades and the second tier blades are fan blades, the former may be made of high temperature resistant material, such as Nimonic (RILML), and the second tier blades may be made of a low density material" such as aluminium alloy. If it is necessary'to provide a heat'resistant material for the root portions of the fan blade titanium could be employed.

The construction and mounting of twotier bladingd'escribed above with reference to a ducted fan gas turbine With the staggering engine is applicable to any other elastic fluid flow machine having concentric and coaxial annular fluid flow paths, through which the working fluid flows in a common axial direction or in opposite axial directions.

Both the fioW paths may constitute flow paths for hot combustion gases and both first and second tier blade-s would then be turbine blades. Alternatively both sets of blades may be compressor, or fan, blades, the radially inner of the two annular fiow paths directing pressure air to a combustion chamber whilst the radially outer flow path provides a stream of pressure air for any other requirement e.g. as by-pass air. Again the first tier blades may be employed as compressor blades and the second tier blades may be used as turbine blades. Materials for the construction of the two tiers of blades are selected in accordance with the function.

It will be understood that the advantageous stress condition outlined above with reference to FIGURE 4 will be obtained with variations of two tier blading mentioned in the preceding paragraph. Arrangements of blading in two tiers in engines where both tiers of blades are turbine blades, or both tiers are fan blades will give a corresponding advantage in reducing the bending on the first tier blades as compared with co-extensive blades if they are mounted in staggered relationship. Similarly, a corresponding reduction in stress in the first tier blades is obtained where the first and second tier blades are of a different kind i.e. turbine blades on first tier fan blades or first tier turbine blades on second tier fan blades as described with reference to FIGURE 1.

With the above described arrangements of two tier blading the second tier blades may be more readily designed as individual components separate from the first tier blades and the shroud structure and of a difierent material. The arrangement moreover has the advantage of limiting the vibration stresses and introduces a substantial amount of vibration damping.

We claim:

1. A gas turbine engine comprising inner, intermediate and outer wall means defining between them inner and outer coaxial annular flow paths; means for supplying air to one of said flow paths; means for supplying combustion gases to the other of said flow paths; a rotor coaxially mounted with respect to said flow paths; and two tier blading carried on said rotor, said blading including a tier of axial flow compressor rotor blades extending across said air flow path, a tier of axial flow turbine rotor blades extending across said combustion gas flow path, there being equal numbers of blades in said two tiers, a circumierentially extending shroud ring connecting the tips of the blades of the inner tier and forming part of said intermediate wall means dividing said flow paths, and means mounting the blades of the outer tier on said shroud at positions circumferentially alternating with the tips of the blades of the inner tier.

2. A gas turbine engine according to claim 1, wherein the shroud ring is made up of a plurality of segments, one carried on the tip of each blade of the inner tier, said segments abutting circumferentially and the abutting portions being formed with root seatings, the blades of the outer tier having roots engaging in said seatings.

3. A gas turbine engine according to claim 1, wherein said outer path is the air flow path and said inner path the combustion gas flow path, said outer and inner tier blades being compressor and turbine blades respectively.

4. A gas turbine engine according to claim 3, wherein said compressor blades are formed of a different, lower density material from said turbine blades.

5. A gas turbine engine according to claim 1, wherein the blades of said two tiers are cambered in opposite senses in relation to the direction of rotation of the rotor,

References Cited in the file of this patent UNITED STATES PATENTS 2,391,623 *Heppner Dec. 25, 1945 2,411,124 Baumann Nov. 12, 1946 2,428,330 Heppner Sept. 30, 1947 2,505,660 Baumann Apr. 25, 1950 2,646,209 Galliot July 21, 1953 2,692,724 McLeod Oct. 26, 1954 FOREIGN PATENTS 189,131 Great Britain Mar. 1, 1923 309,181 Great Britain Dec. 12, 1929 

